The Cost of Space Transportation. The cost of transporting a payload to a low earth orbit can be assessed in terms of the energy required to place a pound in orbit. A pound in orbit has the kinetic energy equal to approximately 3.8 kilowatt hours per pound. When aerodynamic drag and other losses encountered in placing a pound in orbit are taken into consideration the energy required rises to approximately 5.25 kilowatt hours per pound. With the electricity costing ten cents a kilowatt hour, a perfectly efficient machine using electricity to place a payload in orbit would do so at a cost of $0.50 per pound
It is this basic calculation regarding energy and cost that has led some people to speculate that low cost space transportation can only be achieved by an elaborate apparatus employing electricity such as electromagnetic cannon or various exotic rotating skyhooks and the like. In fact ordinary rocket technology is relatively efficient. For instance the Saturn V launch vehicle which was used to send men to the moon placed one pound in orbit for every 22 pounds of fuel used. The majority of fuel, or more precisely propellants, utilized consisted of liquid oxygen--an industrial commodity costing perhaps five cents a pound. The other fuels utilized by the Saturn V were RP1--a special grade of kerosene--and liquid hydrogen. RP1 is relatively expensive, perhaps a dollar or more a pound, however similar propellants such as liquified propane and liquified natural gas which have better performance cost between 10 and 20 cents a pound. Liquid hydrogen costs $1 to $3 a pound.
The cost of the fuel for a reasonably efficient launch vehicle such as the Saturn V amounts to between three and five dollars per pound placed in orbit. This simple calculation shows the rocket to be a relatively efficient transportation device.
In other transportation systems, for example airlines, the cost of transport may be on the order of two and a half times the fuel cost. This would place the cost of placing a pound in orbit at approximately $10 a pound. On the other hand, in operating the family car the cost of the fuel is about three cents a mile or perhaps one tenth the overall cost of transportation by car. If these standards were to apply to launch vehicles the cost of transporting a pound to orbit might be expected to be somewhat less then $30 a pound. In fact, existing transportation systems, at least in the Western world, cost between five and ten thousand dollars to place a pound in orbit. Accurate costs for Russian and Chinese launch systems are not available but are speculated to be several times less costly than European, Japanese, or United States launch systems. Nevertheless, there is a disparity factor of between 100 and 1,000 between the actual cost of space launch and what it might be expected to cost on the basis of physics and economic comparisons with other transportation systems.
One of the underlying reasons why space transportation has so far failed to drop sufficiently in cost over time relates to the so-called mission model or market for space transportation. Up to this time all space transportation systems have been developed by governments. Governments need a sound basis for justifying the expenditure of public funds. Thus, rather than look at the prospective increased market for space transportation as the cost falls, governmental agencies such as NASA and the Air Force have looked to the currently manifested payloads and some reasonable projection therefrom to determine the potential size of the space launch capability which can be reasonably justified.
Current United States space launch capability is equivalent to approximately 600,000 pounds a year placed into low earth orbit. Rarely have mission models projected a requirement of more than 2 million pounds a year into orbit. This space transportation model combined with a preference for launching large payloads has meant that most government studies considering new space transportation systems have focused on maximum launch rates of 40 to 80 launches a year.
At the same time government planners of launch systems typically must consider five to twenty launches as more likely to be flown and therefore typically make the trade-off between development costs and recurring flight costs based on this lower range of five to twenty flights a year. With the rise of systems engineering, space system design has been managed and planed within a system which dictates that space systems must be built to a set of predetermined requirements which in turn are derived from a combination of political considerations, mission models, and budgetary constraints.
By the time the requirements have been set and design of the vehicle is handed over to the vehicle designers, any designs capable of achieving low launch costs are invariably inconsistent with the requirements and so are not presented by designers for consideration. As an example of this dilemma consider a fully reusable launch vehicle capable of putting 20,000 pounds into low earth orbit and having a turn-around time of 48 hours. If a very small fleet of this vehicle (small, that is, in comparison to any other transportation system), say 20 vehicles, were built and flown at design capability, the yearly payload transported to low earth orbit would be approximately 80 million pounds--forty times higher than even the more optimistic mission models put forth by government and industry planners.
A modern airliner such as the Boeing 777 has a development cost of two or three billion dollars. If only four airframes were built and if each plane made two flights a year the true cost of a coast to coast flight for one person could easily exceed a quarter of a million dollars, not counting the facility costs involved in the airports, the air traffic control systems, the supply and maintenance systems, etc. Thus one part of the problem of developing low cost space transportation has been improper choice of the mission model which has led to improper criteria for the vehicle design.
The other difficulty is that no known vehicle design has been demonstrated to be capable of achieving routine reusable operation within the current state of the art of available space systems. The one exception to this may be the mass production of an expendable launch vehicle which logically would result in some fairly significant overall reduction in cost. This appears to be what the Russian launch capability has achieved and accounts for their significantly lower launch costs as opposed to United States, Europe, and Japan. However the degree to which costs may be reduced through this approach is limited.
One proposed approach to providing a cost-effective launch vehicle is to build a vehicle like an airplane on the basis that if it looked like an airplane it may cost like an airplane. From such a philosophy vehicles such as the NASP (National Aerospace Plane) and HOTOL have arisen. However these systems face fundamental technical challenges. Hypersonic air-breathing flight beyond Mach 3 or 4 has not been demonstrated. Nor have necessary subsystems such as the propulsion system, the heat shield system, or the tankage system. Thus such systems as the NASP, while perhaps technically possible, represent tremendously costly investments with the very real possibility of never achieving success. To date over two billion dollars have been spent on the NASP concept without leading to even the production of a demonstration vehicle.
Another low cost launch system concept is the use of a single-stage-to-orbit rocket. Achieving orbit with a single stage is a feasible and readily demonstrated proposition. Many past or existing stages can be shown to be capable of placing a payload in low earth orbit. Examples are the Atlas vehicle which if it does not stage its boost engines can marginally achieve orbit with zero payload. Other examples are the second and third stages of the Saturn V launch vehicle which, if they were reconfigured for ground launch, each would be capable of putting significant payloads into low earth orbit. The shuttle external tank, if integrated with the shuttle liquid propulsion system, can also be utilized to demonstrate the feasibility of a single-stage-to-orbit vehicle.
The ability of a rocket to achieve a given velocity is governed by the rocket equation. If the total velocity through which the payload must be accelerated is known and the average or effective ISP of the stage is known the mass ratio of the vehicle may be readily determined. The mass ratio is the initial or gross weight of the vehicle divided by the burnout or final weight. The total velocity required to reach low earth orbit is a combination of the orbital velocity of approximately 25,500 feet per second and various velocity losses which are incurred in a actual launch. These including losses due to drag, gravity and losses due to potential energy gained. The component of earth's rotational velocity which is aligned with the direction of launch is approximately 1 400 ft/sec and reduces the total velocity required to achieve orbit. Thus the total characteristic velocity is often taken to be between 29 000 and 30,000 feet per second for an eastward 28 degree latitude launch. A state of the art average-to-high performance launch vehicle using hydrogen engines would require a mass ratio to achieve low earth orbit of between eight and nine, meaning that between twelve and a half and eleven percent of the launch weight can be placed in orbit. Existing vehicle stages, particularly the Saturn V second and third stage have structural fractions on the order of nine or ten percent.
So such stages should be able to place between one and three percent of their gross weight, into low earth orbit as payload. While seemingly a minute fraction of the gross weight one to three percent is quite respectable compared to existing vehicles, with the shuttle achieving perhaps one percent and the Saturn 5 achieving somewhat over four percent. This would appear to demonstrate the feasibility of single stage to orbit transportation and does so, but only for expendable vehicles. A reusable vehicle by almost any realistic assessment requires an increase in the dry weight of the stage by approximately 30 percent to account for the heat shield and recovery system. Thus existing single stage to orbit designs generally fall into one of two classes. The vehicle designers assume an improvement in vehicle technology, in light weight structures, tanks, subsystems, and assume engines of improved performance and thrust to weight ratios. These vehicles suffer from problems similar to the air breathing concepts. As such as they have very high development costs without any assurance of eventual technical success. The other approach in developing single stage concepts is to assume less dramatic improvements in technology and so design a vehicle with a large gross takeoff weight in relation to the payload. Thus the payload fraction of this design choice is often less than one percent of the gross weight of the vehicle and less than ten percent the weight of the vehicle structure. This makes the vehicle extremely sensitive to small increases in the weight of the vehicle structure. A small overall increase in vehicle structure and subsystem weight results in no payload at all. The very small margin for weight growth results in an extremely risky development program in that a state of the art vehicle might be built and in the final analysis to have no payload capability.
The most logical approach would appear to be a two stage vehicle. And initially a two stage vehicle was proposed to follow the Apollo era and introduce reusable space transportation. However two stage vehicles are thought to have high development costs in that two optimized stages must be developed and built. Further because the first stage does not achieve sufficient velocity to circle the earth and so return to the launch site the first stage will typically, depending on vehicle design, reenter two to six hundred miles down range of the launch site. If the vehicle is launched over land, the recovery of the first stage is difficult. If the vehicle is to be launched over a range of orbit inclinations, landing the first stage and transporting the stage back to the recovery site is especially difficult. If the launch is conducted over water the first stage returns to a water landing where it is invariably contaminated by salt water. The salt water recovery can necessitate extensive refurbishment and at the same time ocean landings carry a small but significant loss associated with sea states and loss of buoyancy in the recovered stage. For a truly low cost system, even one vehicle lost in a hundred flights is a significant increase in overall operating cost.
The result of this analysis has in the past been an attempt to design a flyback first stage through the use of wings. This is a difficult design task as the empty stage has most of its mass concentrated in the engines and thrust structure which are located in the rear of the stage, so it is difficult to achieve a stable flying platform without significant weight penalties. Additional weight penalties are incurred if an air breathing propulsion system is carried on board. However without the air breathing system, risk of vehicle loss is increased. In either case facility costs are extensive in that wide long runways must be provided at every vehicle launch site. One concept which has been suggested is the so called popback booster where the first stage has sufficient propellent remaining after staging to flyback under rocket power to the launch site. This system while possibly feasible incurs a large penalty in performance and an increased complexity as the rocket propulsion system must be restarted after staging and must be capable of the significantly lower thrusts required for the popback maneuver. Another concept disclosed by U.S. Pat. No. 3,285,175 utilizes a first stage powered by air breathing turbo-rocket engines or other type of air breathing engine. The air breathing stage with an upper stage mounted thereto is launched vertically and recovered vertically at the launch site. This system, while providing theoretical advantages, has practical difficulties. Air breathing engines have thrusts to weight ratios in the range of six or less especially as flight velocity increase. The vehicle can not maintain thrust as the vehicle exits the atmosphere and this presents control or staging problems. Still another concept is that developed by E. Sanger and I. Bredt and later advocated by Philip Bono of launching the vehicle from a captive first stage which runs on track along the ground or up a mountainside. Such systems while improving on the performance of a single stage vehicle, are only capable of reducing velocity required by the orbital stage by little more than a thousand feet per second. Such concepts are often applied to the second type of single stage vehicle, one using near term technology and having very little payload fraction to marginally improve the payload fraction. However the cost of doing this is significantly increased vehicle design complexity and facility costs for very marginal improvement in vehicle performance. Another approach, which has been adopted by the shuttle, the Ariane V, Energia, and the H-2, is to employ a zero stage or half stage which consists of solid or liquid strap on boosters which burn in parallel with what is essentially a single stage vehicle thereby improving its performance sufficiently to achieve a reasonable payload fraction. If these systems were used in a fully reusable system they would have most of the problems associated with a two stage vehicle-that is, recovery of the first or aero stage down range. What is needed is a launch vehicle system which is fully reusable, demonstrably feasible with current technology having adequate performance margins, and which returns all stages to the launch site.